1. Field of the Invention
This invention relates to supersonic flight vehicles and more particularly to a supersonic engine variable geometry inlet for optimizing engine performance.
2. Description of Related Art
Aircraft engines which are designed to operate at speeds ranging from takeoff, through subsonic and transonic, and into the supersonic regime, require complex air inlet configurations in order to operate efficiently throughout the entire operating range. At lower subsonic speeds, particularly at takeoff, it is desirable to allow the engine maximum access to air, since at these lower speeds there is no substantial "ramming" effect produced, whereby air is literally forced into the engine. Hence, in the low end of its operating range the engine must depend on its ability to draw air in to satisfy its large air demand. The amount of air the engine is able to draw in is a function of the area of the throat, i.e., the point along the length of the inlet at which the main airflow passageway is most constricted. In general, the larger the area of the inlet throat, the greater the amount of air the engine can draw in. At transonic speeds, the airflow demands of the engine may also supersede the efficient supplying ability of the inlet since the inlet throat becomes choked. Thus, it is desirable to be able to enlarge the minimum total cross-sectional area of the inlet passageway in order to satisfy engine transonic airflow demand.
It also is well known in the art that efficient supersonic operation of the engine requires a much smaller throat than at lower flight Mach numbers. Furthermore, it is well known that the inlet must be "started," i.e., that the internal airflow in the main airflow passageway of a diffuser having the inlet be changed from subsonic to supersonic. It is known to start the inlet by translating the inlet centerbody and the throat to provide maximum airflow through the inlet as the aircraft speed increases to supersonic speeds. The design of mixed-compression supersonic inlet systems for long-range supersonic cruise aircraft has centered mainly about the design Mach number requirements. Off-design requirements have generally taken secondary roles in the establishment of final designs. As a result, the off-design engine demand airflow matching and performance requirements have been satisfied by complex variable geometry systems, which, in some cases, may not assure efficient inlet to engine matching compatibility.
It is well known in the prior art to provide supersonic and hypersonic aircraft with variable geometry inlets for the efficient matching of engine airflow requirements over a wide range of flight Mach numbers. At the higher supersonic flight Mach numbers where a combination of external and internal compression is required in order to extract high performance from the captured airflow, the inlet throat area must also be varied. Typically this is accomplished by axially translating one compression surface, typically on an inlet centerbody wedge for two-dimensional inlets and on a conical centerbody for an axisymmetric inlet, relative to the another compression surface either fixed or variable. When this is done typically both the captured inlet airflow and the inlet throat area are simultaneously varied, i.e. independent control of airflow and throat area is not achieved. Such an inlet is described in U.S. Pat. No. 4,007,891 entitled "Jet Engine Intake System".
Also typically employed is the modulation of the supersonic compression turn angles to vary the strength of the oblique shockwave while also varying the inlet throat area. These standard approaches can produce inlets providing high performance levels at a selected design flight Mach number. Their performance at other flight Mach numbers ("off-design"), however, is typically less than it could have been had the inlet been designed for that particular flight Mach number. The reason for this is that although utilizing significant geometry variation, these illustrated standard approaches do not provide enough degrees of variable geometry freedom to closely approximate the optimum inlet flowpath at each flight Mach number. Typically, conventional inlets are unable to de-couple airflow capturing capability from throat area variation capability.
Conventional two-dimensional inlets with or without variable geometry capability operating at high supersonic speeds will produce a complex shockwave flow field having a shock from the leading edge of the centerbody or wedge intersecting a shock from the leading edge of the cowl within the inlet. At the intersection of the two shocks, unless they are balanced, a slip plane will form extending downstream from the shock intersection causing an expansion wave or shock to fall on the compression ramps of the wedge centerbody. This reduces the compression ramp's effectiveness for performing the high pressure recovery function for which the supersonic diffuser inlet is designed and can be a significant source of flow profile distortion for the engine.
What is desired therefore, is an inlet geometry which produces a balanced shockwave system at all supersonic flight Mach numbers and for which the last cowl shockwave always impinges at the throat of the inlet. An important advantage of the impinging cowl shockwave always intersecting the center wedge throat, is simplification and possibly increased efficiency of the center wedge boundary layer bleed system.
It is well known to use center wedge boundary layer bleed systems on the wetted surfaces of the compression ramps of the wedge. However, when inlet shocks impinge at different axial positions on the center wedge, the region of surface boundary layer bleed must be enlarged proportionally. Furthermore, this requires, in many cases, several boundary layer receiving plenums on the back side of the compression ramps to accommodate the pressure rises across each of these shocks. They must be compartmentalized and pressure isolated from each other to prevent boundary layer separation on the supersonic compression side of the ramp which results from recirculation zones in the subsonic boundary layer.
There is a great and long felt need among supersonic aircraft and aircraft engine designers for a two-dimensional supersonic inlet that can operate efficiently at both design and off design Mach numbers with a minimum amount of drag and minimum distortion of the flow profile into the engine. The present invention is directed towards a two-dimensional variable geometry inlet effective to operate at supersonic speeds efficiently with a minimum amount of drag and with low values of flow profile distortion into the engine.